Close Air Support Aircraft

Brian Smith

December 11, 2014

Abstract

This report lays out the process and results in designing a new aircraft for the close air supportmission, such that it can replace the currently aging aircraft designs used for this purpose today. The designprocess, as outlined in the following sections, began with rough estimations of the intended capabilities of theaircraft and continued into more detailed tuning of the component sizes, characteristics and layout. For thisaircraft, the new design was intended to have a combination of capabilities from its two main predecessors(one American and one Russian-designed), such that it could have the strengths of both aircraft dependingon the portion of the mission. Specifically, this aircraft was intended to have the speed of the Russian Su-25,while also having the capability to contribute as much to the fight on the ground as the American A-10A.

In order to meet these requirements, the new design was placed in between these two aircraft interms of weight, strength and payload capabilities. In order to increase its speed (which could help savelives on the ground and the pilot’s), the aircraft was given more powerful engines than either of the previousdesigns. In order to confirm the stability of the aircraft, final changes were made to the location and sizeof the wings and the horizontal and vertical tail sections. Finally, estimations were made for the potentialcosts of researching, designing, and producing this new aircraft.

At the conclusion of the design process, it was determined that this new aircraft design was successfulin improving on the aspects of of its predecessors such that it nearly matched the speed capabilities of thefaster Su-25 while maintaining most of the strength, redundancy and other characteristics that would allowit to perform the close air support mission with as much success as the A-10. In addition, the new designproved to be comparable in price such that it could potentially replace these older aircraft in the future andbecome a financially successful aircraft as it fills a role that could gradually become vacant.

1 Initial Design Proposal

I propose to design a new aircraft for the purpose of providing close air support (CAS) capabilities for themilitary of the future. CAS aircraft are designed to support friendly ground or naval forces, and thus arerequired to fly low and slow enough to contribute to the effort. They are also designed to be able to takeoff in relatively short distances, to survive flight through enemy territory at low altitude and withstand thepossibly substantial damage that could result. With the current CAS aircraft in use throughout the worldaging, as most of them were originally designed in the 1970s or early 1980s, there will soon be a need fora new aircraft to fill that role. This includes the Fairchild Republic A-10A, for which retirement has beenbrought up many times in the last two decades, yet has been kept in service due to its singular design toperform that mission. Although newer or improved technologies such as unmanned aerial vehicles (UAVs)and guided missiles or bombs have been able to perform similar duties for precise ground targeting, theexcellent coordination with ground forces that CAS requires indicates that there could still be a need for anew manned aircraft to be able to fly close to the action.

The goals for this new aircraft design are to improve the performance capabilities of current CASaircraft while maintaining a similar amount of ordinance payload, redundancy and resistance to damage thatsuch missions require. Maximum speed and cruise speed are where potentially crucial improvements couldbe made, as increasing these would lead to a faster response to the need of ground troops and an improved

1

ability to escape from danger in the combat zone. The maximum speed could potentially be increasedto high subsonic or even supersonic speeds, with a cruise speed near or over 500 mph. Hopefully thesetypes of improvements could be made by lowering the empty weight of the aircraft and increasing availablethrust, using newer materials for construction and more modern engine designs to increase the aircraft’sfuel efficiency and thrust-to-weight ratio. The intended maximum takeoff weight would be less than 40,000pounds, a majority of weight reduction coming from the empty weight of the aircraft. Other characteristicssuch as combat radius and takeoff distance would be kept similar to those of current aircraft or slightlyimproved if possible.

Previous CAS aircraft designs include the currently-used A-10A and the Sukhoi Su-25, along withprototype designs such as the Northrop YA-9A and the Ilyushin Il-102. The current CAS aircraft for theUnited States, the A-10A, is a larger and more powerful design than the Russian Su-25, enabling it to carrymore ordinances and withstand more damage. However, its counterpart has the advantage in speed andcombat radius. Both aircraft are shown in Figure 1 below. The design drivers for the new aircraft wouldbe to achieve the speed range of the Russian design without sacrificing much of the survivability and otherCAS capabilities the American design has. In this way, the new aircraft could better fill the role these hadbeen built for since it possesses many of the abilities of both designs, made possible by newer developmentsin structural materials and engine efficiency.

(a) Fairchild-Republic A-10A (b) Sukhoi Su-25

Figure 1: Comparison CAS Aircraft.

2 Preliminary Estimate of Take-Off Weight

With this design proposal, the primary goal is to design a CAS aircraft with the speed of the Su-25 orbetter while keeping as much of the payload capacity and survivability of the A-10 as possible. As such,the target take-off weight (TOW) was 40,000 pounds, and speed was kept as the most solidified input intothe spreadsheet used to estimate take-off weight. The maximum Mach number was set to be 0.8, with acruise Mach number of about 0.67 which would be higher than that for both the A-10 and Su-25. The cruisealtitude was set to 20,000 feet, similar to that for the A-10 and higher than for the Su-25, so that the aircraftwould be capable of cruising to the target over enemy or contested territory that might have a hard deckaltitude to avoid ground-to-air fire. The operating radius was set to 500 nautical miles so that it wouldhave a larger range than its A-10 predecessor. While range was not one of the main design drivers, it wasdetermined that it could still be brought up closer to the level of the Su-25.

At this point the fuel consumption rates for the engines were not well known, so estimations weremade for the minimum and maximum TSFC, using a listed typical range for aircraft of this type. In Corke’sDesign of Aircraft [1], a general range of 0.5 to 1.2 was given. However, using computed data from theGeneral Electric TF34 turbofan engine of the A-10 [2], a TSFC of 0.4 was chosen for the minimum amount,and 1.0 was estimated for the maximum. The thrust was set to 20,000 pounds to get close to the targetthrust-to-weight ratio of 0.5.

To maintain its ability to contribute to the fighting on the ground, the combat and loiter time forthis design were kept close to those for the A-10, with 15 minutes of combat time and at least 90 minutes ofloiter time. Both were to occur at or under 5,000 feet of altitude. Finally, the structure factor was estimatedusing historical data for smaller military aircraft. The middle range for typical military fighter jets was

2

seen to be from 0.45 to 0.5 [1], so the estimation of 0.45 was made since part of the construction might usecomposites or other lighter materials.

In terms of payload, the design would not be able to have the same amount of maximum capacityas the A-10 and reach a take-off weight near 40,000 pounds. However, it was a goal to keep as much aspossible and to carry more than the Su-25 could. Therefore, the amount of expendable payload was set tobe 10,000 pounds. While this is a fair amount below the A-10’s maximum capacity of 16,000 pounds [3], it isimportant to note that the A-10 is never loaded to full capacity for CAS missions due to a desire to maintainmore maneuverability in the combat zone. Therefore, this design’s 10,000 pounds should be enough to fulfillmost types of missions, and is still more than the Su-25’s maximum capability of less than 8,800 pounds [4].

After taking all of these values into consideration, it was calculated that the maximum take-offweight would be approximately 40,006 pounds. This is very near the initial target weight of 40,000 pounds,and allows the aircraft to reach the speeds of the Su-25 while improving on the range and maintaining muchof the payload capacity of the A-10.

3 Wing Loading Selection

In determining the wing loading for the aircraft, the optimum value for cruise was first calculated using cruiseparameters previously defined in the take-off weight spreadsheet. Using the given aspect ratio, altitude, Machnumber and weight at the beginning of cruise, the wing loading was calculated to be about (W/S) = 85.8pounds per square foot. Also used in this calculation was the coefficient of drag at zero lift (CD,0), whichwas estimated using NACA lift and drag coefficient data for the listed airfoil used by the A-10. Using thiscalculated wing loading value, the area for the wings was found to be about 440 square feet.

The take-off and landing distances were found using this wing area along with the estimated airportaltitude, maximum lift coefficient, thrust capacity, and weight at those two points. The maximum liftcoefficient was estimated using the same NACA data previously mentioned, and the weights at take-off andlanding were taken from calculations of the previous spreadsheet. The take-off distance was estimated to beabout 2,540 feet, while the landing distance was estimated at about 3,020 feet. The validity of this take-offdistance could be verified by the comparison of the wing loading to those of the A-10 and Su-25. Since thewing loading at take-off (90.4 pounds per square foot) is less than those for the other two aircraft, it wouldbe expected that the required take-off distance would be slightly less than the reported values for the A-10[3] and Su-25 [4] of about 2,760 feet and 3,200 feet, respectively.

The instantaneous and sustained turn rates are important in the analysis of the aircraft’s CAScapabilities, as they measure maneuverability in the combat zone. Assuming an altitude of 5,000 feet (dueto its CAS mission that requires close proximity to ground targets), a Mach number of 0.5 or less and amaximum lift coefficient (CL,max) of 1.2, the load factor was calculated to be about n = 4.3. Using this loadfactor and the wing loading value from cruise conditions, the sustained turn rate was estimated to be about9.9 deg/s. This is similar to that for the A-10, calculated from a reported 180-degree turn in 17 seconds,or 10.6 deg/s. While the take-off distance was improved from those of the A-10 and Su-25, this aircraftmaintains near the same amount of maneuverability of its A-10 predecessor.

4 Wing Design

In designing the wing, the first input values were the conditions for cruise and the wing area as found fromprevious portions of the aircraft design process. With the cruise speed at M = 0.67 and a combat speednear M = 0.5, historical data (from Figure 4.12 in [1]) suggested the leading edge sweep angle should bebetween Λ = 5◦ and Λ = 15◦. As such, the sweep angle was chosen to be Λ = 10◦. Using this sweep angle, itwas seen that in order to minimize lift-induced drag, the taper ratio should be set to λ = 0.4 or less. Sincethe penalty for not pursing a smaller taper ratio was small, the value decided on was the most structurallyreasonable value of λ = 0.4.

Starting with the data found for the A-10’s NACA 6716 airfoil shape [3], the values regarding thecoefficients of lift and drag were estimated. Due to this aircraft’s higher intended speed, the wing’s cross-sectional shape was changed to the NACA 6713 so that the maximum thickness-to-chord ratio would provide

3

for a higher critical Mach number (as seen in Figure 4.4 of [1]). This airfoil has similar lift characteristicsas the NACA 6716, but with a smaller amount of drag at small angles of attack. In fact, this airfoil shapematches that for the A-10 at its wing tips.

After going through an iterative process using the calculations from the spreadsheet, it was deter-mined that slightly increasing the aspect ratio (AR) from the original intended value of 5.5 to 6.0 made asignificant reduction in the drag on the wing. With the new value of AR = 6.0, the wingspan of the aircraftwas calculated to be about 51 feet, making it 6 feet less than that for the A-10, with a chord length of 12.2feet at the root and 4.9 feet at the tip. The total drag force on the wings was calculated to be 1,790 pounds.The plan view for one wing is shown below in Figure 2.

Figure 2: Wing plan view for design aircraft.

5 Fuselage Design

For designing the fuselage size, the cruise conditions were required as inputs for the spreadsheet calculations,with a cruise speed of M = 0.67 and cruise altitude of 20,000 feet. Since this aircraft is only designed to carryone pilot, there was no need to consider passenger or cargo volume that would be required for commercialaircraft. Thus the primary volume requirements for this aircraft are for the single pilot crew, the controlsystems, and what is necessary to complete the CAS mission. By designing for the weapon hardpoints toall be mounted underneath the wings, the aircraft can be re-armed more quickly and the largest contributorto required fuselage volume is fuel. With a calculated fuel weight of about 11,000 pounds (from the originaltake-off weight spreadsheet), the required fuel volume was calculated to be about 240 cubic feet using anestimated specific volume for the aviation fuel. Self-sealing bladder-type tanks are to be used for fuel storage,as this improves the survivability of the aircraft in the event of damage sustained from enemy fire. However,since these tanks cannot store their full volume of fuel, it was estimated that about 300 cubic feet of volumewould be required for the fuel tanks. These tanks would be located in the inboard portions of the wings aswell as in the fuselage above the wings. Locating the tanks there also improves the aircraft’s survivability,as being above the wings protects the fuel tanks in the fuselage from ground fire.

Since the aircraft is never intended to fly at supersonic speeds, a fuselage shape for minimizing wavedrag would not be necessary. Therefore, at this preliminary stage a simple cone-cylinder shape was used forthe fuselage. Using the previously-mentioned volume requirements and comparisons to the currently-usedCAS aircraft, the maximum diameter for the fuselage was estimated to be approximately 5 feet. With anintended length under 50 feet (using a similar decrease in size from the A-10 to that for the wingspan), afineness ratio (fuselage length over diameter) was chosen so that (l/d) = 9, giving the aircraft a length of 45

4

feet. This falls within the typical range of 8 ≤ (l/d) ≤ 12.5 for subsonic aircraft, and while these values aresubstantially larger than the ratio for minimized drag of about (l/d) = 3.3, such a value would be impracticalfor the aircraft’s designed mission and the penalty for having a larger ratio is acceptable.

Using all of these inputs, the viscous drag force on the fuselage was calculated to be almost 460pounds. Since the aircraft flies at only subsonic speeds, there is little to no wave drag, so the total drag onthe fuselage is equal to that for the viscous drag. This total drag on the fuselage is noticeably less than thatfor the wings calculated in the previous section of 1,790 pounds. Adding these two values together, it wascalculated that the total drag on the fuselage and wings is approximately 2,250 pounds. As can be seen inFigure 3, this aircraft design is generally similar in fuselage and wing proportions to the Su-25, while beingslightly smaller than the A-10.

Figure 3: Top view of fuselage and wings for A-10, design aircraft, and Su-25 (left to right).

6 Horizontal & Vertical Tail Design

The first step of the tail design process was selecting a tail configuration to best meet the requirements of theClose Air Support mission. Using the A-10 and Su-25’s tail arrangements as the first two for comparison,the inverted Y-tail was chosen as the third tail design to be considered. It was determined through closeconsideration of the type of mission the aircraft must complete that an H-tail arrangement similar to thatfor the A-10 would be the most suitable design. Putting the vertical stabilizers at the ends of the horizontalstabilizers gives the double benefit of them serving as winglets for the horizontal stabilizer (reducing therequired size), and increases their contribution in spin recovery. In addition, having two vertical stabilizersallows for each to be smaller, or for them to be made redundant in case one is damaged from enemy fire.Finally, the shape of the H-tail helps to hide the heat signature of the engines from below and from the sides,making it more difficult for heat-seeking missiles to properly track the aircraft. The only major disadvantageof this tail design is the extra weight necessary for the horizontal stabilizers to be able to support the verticalstabilizers, though this amount of weight could be afforded for an aircraft nearly 5 tons lighter than its A-10predecessor.

In order to complement the H-tail configuration, it was determined that the best position for theengines would be on top of the fuselage in front of the tail. Using this and the specific coefficients thatcorresponded to an H-tail arrangement, it was found that the horizontal and vertical stabilizers should bea distance from the wing of just under half of the entire length of the fuselage. This coefficient was used tomultiply the tail coefficients by in order to size the horizontal and vertical tails. Using values in betweenthose for a combat fighter and combat trainer, the final coefficients for tail sizing were calculated to beCV T = 0.035 and CHT = 0.57. Finally, the horizontal and vertical stabilizer areas were calculated to beSV T = 37 square feet and SHT = 111 square feet.

The cross-section shape for the tail was chosen to be that of a NACA 0012 airfoil. The thickness ofthis symmetric airfoil was less than that for the main wing, leading to a slightly smaller coefficient of dragfor the airfoil shape while maintaining enough thickness for the horizontal tail to be able to support thevertical tails. The horizontal tail was given a leading edge sweep angle of ΛLE = 15◦ (slightly larger than

5

that for the wings of 10◦), an aspect ratio of AR = 3.5 and a taper ratio of λ = 0.4, both in the middle tohigher range for combat aircraft. The vertical tail was given a sweep angle of ΛLE = 40◦ (in the lower partof the historical range), an aspect ratio of AR = 1.4 and a taper ratio of λ = 0.4, which were again in theupper end of the ranges for most combat aircraft. Using all of these input values, the viscous drag on thehorizontal and vertical stabilizers could be calculated. The drag on the horizontal tail was calculated to beDHT ≈ 368 pounds, while the drag on the vertical tail was DV T ≈ 129 pounds.

A comparison of top view of the design aircraft (with the newly-added tail sections) to those of theA-10 and Su-25 is included below in Figure 4. A side view of the design aircraft is shown in Figure 5.

Figure 4: Top view with tail sections for A-10, design aircraft, and Su-25 (left to right).

Figure 5: Side view for design aircraft with tail sections.

7 Engine Selection & Design

The first step of the engine selection process was completed in the previous step of the total design processwhen determining the tail design. It was determined then that the engines of the aircraft would be locatedon top of the fuselage between the wings and tail. This location both protects the engines from ground firemore than locating them under the wings, and helps to hide the heat signature of the engines to protect theaircraft from heat-seeking missiles. In order to maintain the survivability of its predecessors, the aircraft wasgiven two engines with the available power to be able to safely operate with one engine damaged or off. Inorder to have this capability, it was determined that the required thrust for a single engine should be ableto match the total drag on the aircraft at cruise conditions, as opposed to only half of it.

The reference engine was chosen to be the A-10’s General Electric TF34 Turbofan engine, rated for9,065 pounds of thrust force. This older design was chosen since few other currently-existing engine designsare rated in the same range of thrust as needed for this aircraft. When comparing the reference engine’sthrust at cruise conditions to the required thrust, it was found that the engine produced more than required.The required engines could thus be smaller than the TF34, giving 8,007 pounds of thrust each. This gavea total thrust of 16,014 pounds, providing a thrust-to-weight ratio of (T/W ) = 0.40. However, one of thedesign drivers for this aircraft was to have a thrust-to-weight ratio of (T/W ) = 0.5, which would require atotal thrust of approximately 20,000 pounds between the two engines. Bringing the required thrust for eachengine up to 10,000 pounds, it was determined that the required engines would be only slightly larger thanthe reference engine. The required engines would each have a dry weight of W = 1, 885.5 pounds, a length

6

of L = 105 inches, a maximum diameter of D = 51.5 inches, and a mass flow rate of m = 84.9 pounds massper second.

A comparison of top view of the new aircraft to those of the A-10 and Su-25 is included below inFigure 6. A side view of the new aircraft is shown in Figure 7, and a front view is shown in Figure 8.

Figure 6: Top view with engines for A-10, design aircraft, and Su-25 (left to right).

Figure 7: Side view for design aircraft with engines.

Figure 8: Front view for design aircraft with engines.

8 Take-Off & Landing Distances

The calculation and estimation of the take-off and landing distances required many input values from previousportions in the design process, from the wing design to the take-off and landing weight conditions. For thismilitary aircraft, the minimum requirements were set to those from MIL-C5011A rules, including minimumrolling friction coefficients (with and without braking) and a 50-foot obstacle height. In addition, the FAR-25requirement that the quoted landing distance be 1.6 times longer than calculated could be removed.

Using the airfoil section and wing planform data from the wing design section, the maximum 3-D lift coefficient CL,max was calculated using an estimation from the “Flaps” spreadsheet. This requiredestimations for the parameter values needed to convert the data for a 2-D airfoil to that for the 3-D wingwith slotted flaps. As such, the maximum lift coefficient was calculated to be CL,max = 2.32. The addeddrag from deflecting the flaps was safely estimated to increase the coefficient of drag by ∆CD,0 = 0.05.

7

An important area of estimation was that for the rolling friction coefficients µ for the landing gearon the ground during the take-off and landing rolling distances. In order to account for multiple types ofrunway surface that flying out airstrips closer to the battlefield as part of the CAS mission might require, acoefficient of friction of µTO = 0.05 was chosen for take-off. This would be high enough account for taking offof concrete, hard turf, or dirt runways as the mission requires. In the opposite sense, a coefficient of frictionfor landing should be low enough to account for stopping on a similar range of surfaces. With braking, arolling friction coefficient of µL = 0.4 was used.

With these values in place, the final take-off distance was calculated to be about sTO = 3, 410 feet,and the landing distance sL = 3, 010 feet. These values were slightly higher than expected earlier in thedesign process, but since these are “safe” and roughly-estimated values, they are acceptable for this aircraftand its mission.

9 Enhanced Lift Design

For this aircraft the take-off and landing distances were not primary design drivers, the effort was still madeto minimize these distances given the aircraft’s geometry and thrust capability. This required enhancing thelift of the wings for take-off and landing to decrease the required speed for both. Using the previously-chosenNACA 6713 airfoil, the attempt was made to maximize the three-dimensional coefficient of lift for the wingswithout adding any active lift enhancement for design simplicity.

Similarly to the two current comparison aircraft, it was decided to use a slotted flap design with arelatively large amount of wing area used for the trailing edge flaps, so that Sf/Sw = 0.40. The flaps werechosen to deflect to a maximum angle of δf = 40◦. The ratio of the chords for the flaps and the total wingwas set to cf/c = 0.25. The value required for the wing to be considered to have a high aspect ratio wascalculated to be AR = 2.90, much less than the actual aspect ratio of AR = 6. Using the accompanyingfigures for slotted flaps, the required coefficient values could be approximated.

After all of the calculations were concluded, it was determined that the three-dimensional wing byitself had a CL,max-value of 1.78. The trailing-edge slotted flaps contributed the lift with an approximatedincrease of ∆CL,max = 0.559. When adding leading-edge slotted flaps (slats), with an estimated contributionof ∆CL,max = 0.4, it was calculated that with all flaps deployed the final three-dimensional coefficient of liftcould be increased to CL,max = 2.496.

10 Structural Loads & Refined Weights

This portion of the design process dealt with the affirmation or adjustment of estimates made in previousportions, as well as confirmation of the plausibility of many design factors. First off, the wing and fuselageload distributions were calculated, along with the design load factor for the aircraft. These were done usingthe “Loads” spreadsheet. Dealing first with the “Wing” sheet, the input parameters for the wing shape aswell as the cruise and weight conditions were used to calculate the lift distribution for the wings. Importantinputs to note include the increase in lift due to flaps ∆Lflap, the fuel weight in the wings, and the absenceof engine weight on the wings due to their location on the fuselage. The increase in lift from the flaps wasestimated using the ratio of CL,max with flaps to CL,max without and the portion of the wingspan with flaps.The fuel in the wings was estimated to be less than half of the total fuel weight, as a majority of the fuelwould either be located in the wingbox or in the fuselage above the wings. Thus the load distribution for thewings was plotted for elliptical and trapezoidal planform shapes, along with the average between the two.

For the fuselage structure analysis on the “Fuselage” sheet, inputs for all of the component weightswere needed, along with their relative position along the fuselage as a percentage of total length from noseto tail. With these input values from previous portions of the design process it was determined that by onlyslightly modifying the fuselage length and center of lift locations from their original estimates, the initialpitching stability margin estimate could be made positive, such that S.M. = 0.166. This meant that theaircraft was found to be statically stable with this initial estimate. In addition, the shear force and bendingmoments could be estimated as a function of distance along the fuselage, and were plotted in the spreadsheetas seen below. The final part of the “loads” spreadsheet estimated load factors for different portions of the

8

aircraft’s flight regime. As expected, the case for the highest angle of attack proved to give the highestrequired load factor with n = 5.08, with instantaneous turn rate as the second highest. This would serve asthe load limit, so that with a safety factor of F.S. = 1.5 the design load factor was found to be n = 7.61.

At this point, the weights of the different components of the aircraft structure were recalculatedusing the “Refwt” spreadsheet. These values were calculated for three different aircraft types: militaryfighter, transport, and general aviation. Since this aircraft would behave most like a military fighter amongthese three, those values were used. In addition, many coefficient values for the calculations varied betweenthe types of aircraft, so the fighter’s values were also used for those. Using these input values from the wing,fuselage and tail design stages, as well as coefficient values for the simple three-wheeled landing gear, thefinal weights for those structures were calculated. Adding on the estimated installed engine weight, it wasfound that the expected weight for the other portions of the aircraft structure was within the accepted range(±5%) of the target weight. This meant that the original estimate was close enough to the final one thatweight reduction or changes to the structure factor were not necessary. The weight distribution for eachcomponent of the aircraft is shown below in Figure 9.

Figure 9: Weight distribution for design aircraft, by component.

Finally, the static margin for static stability could be recalculated using the new component weightvalues and the first portion of the “Stab” spreadsheet. By simply entering the new weight values fromthe “Refwt” spreadsheet and their distributed location along the fuselage length from the “Fuselage” pageof the “Loads” spreadsheet, the total magnitude of the weight and moment around the center of lift werecalculated. Using the relative locations of the center of lift and the center of gravity, the static margin wasrecalculated to be approximately S.M. = 0.3. This reaffirms that the intended aircraft design is staticallystable, though this new value was found to be almost twice that of the initial estimate. This proves to bebeneficial for the aircraft, as the aircraft would be able to remain stable throughout the flight, after fuel andpayload with weight located relatively far forward on the aircraft were spent.

11 Static Stability

Previously, the refined weight values were used to calculate the static margin as an approximation of thelongitudinal stability of the aircraft. By moving along through the “Stab” spreadsheet, this stability couldbe confirmed by calculating the longitudinal stability coefficient. By entering parameter values for the wings,horizontal tail and engines, the components’ respective contributions to the total coefficient were found. Thewing effect was found to be stable, while those for the inlet and horizontal tail were found to be unstable.Together, the total longitudinal stability coefficient ended up having a value of CM,α = −2.824. Since thisvalue was negative, the aircraft was found to be longitudinally stable.

9

In order to compute the directional stability coefficient, contributions from the wings, fuselage andvertical tail were found. The computation of the wing contribution required the coefficient of lift for cruiseconditions, the value for which was taken from the “Wing” spreadsheet calculations. The wing and verticaltail effect were found to be stable, while the fuselage effect was found to be unstable as expected. Thesevalues were summed to reach the total directional stability coefficient of Cn,β = 0.0564, such that with apositive value the aircraft was found to be directionally stable. A first approximation for the lateral stabilitycoefficient could be made by simply setting it opposite and equal to the directional stability coefficient. Sincea negative value for this coefficient is required for stability, this coefficient’s estimated value of CL,β = −0.0564means the aircraft is also laterally stable.

In determining the rudder size required for sufficient stability and control in adverse flight conditions,typical parameter values such as a rudder deflection of δr = 20◦ and a potential sideslip angle of β = 11.5◦

were used. The required directional control from the rudder was estimated for two cases: asymmetric thrustand cross-wind conditions. The asymmetric thrust condition accounted for one engine failing, leading to therest of the availably thrust only acting on one side of the aircraft. The cross-wind condition accounted forwinds affecting the overall approach vector for air passing over the aircraft, using the previously-mentionedsideslip angle value. These conditions have the largest effect at take-off or landing, respectively. Calculatingthe directional control Cn,δr for both cases, the one with the larger value would have a larger effect. Thereforethat value was used to estimate the required rudder size relative to the vertical tail area. The cross-windcase was found to have a larger Cn,δr-value, and was used to estimate the minimum ratio of rudder sizeto vertical tail size as about Cr/Cvt = 8%. This value may seem relatively small, but it only serves as aminimum value required to provide stability for these cases.

12 Cost Estimate

In estimating the costs involved in developing, testing, and producing this aircraft design, the spreadsheettitled “Cost” was used. The inputs required for these costs to be calculated included the empty structureweight of the aircraft, its maximum velocity, and the number of aircraft and rate of their production. Also,using the associated costs from the cost-estimating relationships (CERs) of 1970 and 1986, the current-daydollar values could be found using CPI estimates to account for inflation since those years. Since in thespreadsheet the original year was set to 1998, the rate of inflation between 1998 and 2014 was required, andwas found to be about 1.46.

For the number of prototype aircraft it was decided to have two, as per typical designs, built at arate of one per year such that the rate of their production would be 0.083 aircraft per month. After enteringthe final input values for the two turbofan engines with about 10,000 pounds of maximum thrust each, thecosts could be calculated for any number of production aircraft. It was decided that the production aircraftcould be built at a rate of five per month, and the best estimate for the total number of aircraft producedwould be in the area of 400 aircraft. This number seemed appropriate with the knowledge that there wereover 700 A-10s and possibly over 1,000 Su-25s were produced, so that with recent limits on military budgetsless would need to be produced.

Regardless of the number of production aircraft, the total research, development, testing and eval-uation (RDT&E) cost was estimated to be about $1.41 billion. For this and all following values, the 1986estimations were used, as this would be closer to the modern values. With 400 production aircraft, thetotal acquisition costs totaled to about $8.1 billion, with an intended profit of about $865 million. Theunit cost for the first 200 aircraft was estimated to be approximately $27.2 million, while for the rest of theproduction aircraft it would be about $20.2 million. These can be compared to $18.9 million for the A-10 [3]and $18.0 million for the Su-25 [4], both adjusted for inflation since their production. It is important to notethat having more aircraft produced would reduce the unit cost, and it calculated that the unit cost wouldactually be less than for both the A-10 and Su-25 when they had equal numbers of aircraft produced. Thisrelationship can be seen in Figure 10 below. Thus it can be seen that the unit cost of this design is readilycomparable to its predecessors.

10

Figure 10: Unit cost comparison versus number of production aircraft.

Conclusion

In comparison to its two predecessors, the A-10 and Su-25, this aircraft leaned toward the side of theRussian design in size and was able to reach similar speed capabilities. This was done while also enabling theaircraft to have many of the specific characteristics of the A-10 that make it so successful in performing theCAS mission, such as payload capacity, loiter and combat time, redundancy, etc. Some of the few changesneeded to be made during the design process included the reshaping and movement of the wings to improveits lift and stability characteristics. First, the aspect ratio was increased from its initial value of AR = 5.5 toa new value of AR = 6.0 such that the wing’s lift and drag characteristics could be improved. The wings werethen shifted back only slightly from the originally estimated location to improve the stability of the aircraftwhen the fuel and payload are spent. The final change to the design led to an unforeseen improvementin the design, as newly-discovered data for the fuel efficiency of the comparison engines (represented by asmaller reported TSFC) allowed for a greater combat range to be reached with the same take-off weight.A side-by-side comparison of many numerical characteristics for the new design and comparison aircraft isshown in Table 1 below. The final appearance of the design is shown in a 3-view diagram in Figure 11.

Table 1: Final comparison of design to current CAS aircraft.

Design A-10 Su-25Take-Off Weight (lbs) 40,000 50,000 38,800Empty Weight (lbs) 18,000 25,000 21,605Max Speed (mph) 565 440 590

Cruise Speed (mph) 470 340 ≈ 460Combat Radius (nm) 500 300 400

Payload (lbs) 10,000 16,000 8,800Take-Off Distance (ft) 3,410* 2,760 ≈ 2,500

Thrust / Weight 0.50 0.36 0.51Unit Cost ($, 2014) 20.2 million 18.9 million 18.0 million

11

Figure 11: Final appearance and dimensions of the new aircraft design.

This new aircraft design could be considered generally successful, as it proved capable of reachingnearly all of its original intended characteristics in such a fashion that very few changes had to be madealong the way to enable its completion. In addition, this aircraft could potentially lie in the same price rangeas its predecessors did (depending on future demand for aircraft of this type), and thus could be financiallysuccessful in replacing them as the close air support aicraft of the future.

References:

[1] Corke, Thomas C. Design of Aircraft. Upper Saddle River, NJ: Prentice Hall, 2003. Print.

[2] Chiu, Ya-tien. ”A Performance Study of a Super-cruise Engine with Isothermal Combustion inside theTurbine.” (n.d.): n. pag. Polytechnic Institute and State University, 9 Dec. 2004. Web. 6 Dec. 2014.<http://scholar.lib.vt.edu/theses/available/etd-12212004-091959/unrestricted/Dissertation.pdf>.

[3] Jenkins, Dennis R. Fairchild-Republic A/OA-10 Warthog. North Branch, MN: Specialty, 1998. Print.

[4] ”Sukhoi Company (JSC) - Airplanes - Military Aircraft - Su-25K - Aircraft Performance.” SukhoiCompany (JSC) - Airplanes - Military Aircraft - Su-25K - Aircraft Performance. N.p., n.d. Web. 10 Dec.2014. <http://www.sukhoi.org/eng/planes/military/su25k/lth/>.

12